1. Field of the Invention
This invention relates to a rocket nozzle assembly capable of implementing in-flight thrust variation in a controlled manner by actuation of thrust control cylinders. This invention is also related to a rocket assembly comprising a rocket nozzle assembly with thrust control cylinders.
2. Description of Related Art
Rocket motors produce thrust by expelling high pressure combustion chamber gas through a nozzle throat and expanding the expelled gas against the nozzle walls.
Rocket motors, especially tactical solid rocket motors, frequently need some form of in-flight thrust management. Generally, thrust is controlled in flight by the use of predesigned boost-sustain thrust solid rocket fuel grains. For example, the propellant grain is often designed to have a high burning surface area during an initial boost phase of the burn and then transition to a lower burning surface area for the subsequent sustain phase (or remainder) of the burn. The propellant burning response for the sustain phase results in decreased motor pressure and, consequently, decreased thrust.
One of the drawbacks of a two-phase (or multi-phase) propellant grain is that the nozzle throat area is usually optimized for performance of only one of the phases, typically the boost phase. This drawback is especially problematic where it is desired to control the amount of thrust produced during the sustain flight phase of a rocket motor. One manner of varying thrust during the sustain phase is to control the effective throat area. Specifically, if the effective throat area of the nozzle is increased, the combustion chamber pressure will decrease, resulting in an attendant drop in the thrust level. On the other hand, if the effective cross-sectional area of the throat is decreased, the pressure in the combustion chamber will increase, resulting in an attendant increase in thrust level. By controlling the motor pressure through active throat area control, the engine can be operated over a larger altitude range in the atmosphere and the thrust can be modulated for optimal performance. Improved performance may result in improved ranges for tactical missiles.
Several approaches for changing the throat area of a rocket nozzle have been proposed and practiced. One of the most common approaches involves the use of a pintle movable along a nozzle axis relative to the nozzle throat, as described in, for example, U.S. Pat. No. 3,948,042 to Beardsley et al. Generally, a pintle is hydraulically moved axially in one direction along the nozzle axis towards the throat region to decrease the size of the throat, and in an opposite axial direction away from the throat region to increase the size of the throat. As the throat size decreases, the internal pressure increases. On the other hand, as the throat size increases, the internal pressure decreases. In this manner, thrust levels may be varied and controlled by axial movement of the pintle. The pintle design provides flexibility by allowing the nozzle area to be varied in flight in accordance with a particular operation profile and, with some designs, allows for the possibility of multiple different throat sizes. However, the conventional pintle design has drawbacks. For example, actuation mechanisms for the pintle are commonly carried inside of the motor case. As a result, the actuation mechanisms decrease the available case volume into which propellant may be located and raise design concerns over thermal protection and integration of the pintle actuation system. Also, a pintle is subjected head-on to the full force of combustion products passing through the nozzle.
A modification to the conventional pintle is described in U.S. Pat. No. 3,907,222, in which a fustro-conical pintle is mounted on a shaft rotatable on an axis which is transverse to the nozzle axis and upstream from the throat section. Rotation of the shaft 180 degrees about its axis moves the frustro-conical pintle into and out of an annulus-forming position. When the pintle is out of the annulus-forming position, the throat section is operable at a normal (large) throat dimension, substantially unaffected by the pintle. On the other hand, in the annulus-forming position, the pintle is rotated closer to the throat, thereby forming an annulus between the pintle and the inner wall of the convergent section. The annulus is smaller in cross-sectional area than the open throat and, as a consequence, internal pressures of the rocket motor are increased when the pintle is in the annulus-forming position. Thus, movement of the pintle into and out of the annulus-forming position allows for dual-mode control over thrust by control of the throat area.
However, the modified pintle design of U.S. Pat. No. 3,908,222 is not without its own drawbacks. Because the pintle is located along the nozzle axis, the pintle carries the full blowout load of the operating pressure and, therefore, must generate high actuation torques. Also, the pintle of this modified design rotates into either a fully open or closed position and is not movable into intermediate positions to permit continuous variable control over the throat area.
In accordance with the principles of this invention, the above-discussed problems of the related art are overcome by the provision of a rocket nozzle assembly comprising a nozzle insert, first and second thrust control cylinders, and at least one thrust control cylinder-rotating subassembly. The nozzle insert structure provides a converging region that converges in cross-section to meet a throat region located aft of the converging region and a diverging region located aft of the throat region and extending radially outward. The converging and diverging regions and throat region are coaxially aligned with each other along a passageway central axis and collectively define a converging/diverging passageway. The first and second thrust control cylinders are rotatable about respective first and second axes, which are arranged transverse to the passageway central axis. The first thrust control cylinder has a first outer surface with at least one first groove extending transverse to the first axis, and the second thrust control cylinder has a second outer surface with at least one second groove extending transverse to the second axis. The first and second axes both lie in a plane that is normal to the passageway central axis and are parallel to and spaced apart from one another.
The thrust control cylinder-rotating subassembly is operatively associated with the first and second thrust control cylinders to rotate the first and second thrust control cylinders about the first and second axes, respectively, relative to the throat region between an open position and at least one throat-reduction position. In the open position, the first and second grooves face each other from diametrically opposite sides of the throat region to maximize the effective cross-sectional throat area at the throat region. In the throat-reduction position, outer surface portions of the first and second thrust control cylinders intersect and partially obstruct the passageway at the throat region to reduce the effective cross-sectional throat area relative to the effective cross-sectional throat area in the open position. By controlling the effective cross-sectional throat area through which combustion products may pass, it is possible to control the amount of thrust generated when the rocket nozzle assembly is operatively engaged with an operating rocket motor.
The thrust control cylinder-rotating subassembly is preferably capable of moving the first and second thrust control cylinders in a controlled manner to any position between the open position and the fully closed position. In this manner, the effective cross-sectional throat area can be proportionally controlled by moving and holding the variable thrust control cylinders at any rotational position between the open and fully closed position. It is possible, however, to design the thrust control cylinder-rotating subassembly to limit rotational movement to two or more discrete positions, e.g., the open position, the throat-reduction position, and optionally one or more partially closed positions therebetween.
This invention is also directed to a rocket assembly comprising a case, at least one propellant, and at least one nozzle assembly.
These and other objects, features, and advantages of the present invention will become apparent from the following detailed description of the invention when taken in conjunction with the accompanying drawings which illustrate, by way of example, the principles of this invention.